(1) Field of the Invention
The present invention relates to the general technical field of automatic or semi-automatic flight control systems, and more particularly to a method of controlling and regulating the deflection angle of a moving tailplane of a particular type of rotorcraft, specifically a hybrid helicopter.
More particularly, the present invention relates to a rotorcraft and to its system for controlling a movable tailplane, with the deflection angle of the tailplane being determined as a function of specific parameters in stabilized flight at a high cruising speed.
(2) Description of Related Art
The term “rotorcraft” is used to designate an aircraft in which lift is provided in full or in part by one (or more) large-diameter propeller(s) having an axis that is substantially vertical, and referred to as a “rotor” or indeed as a “rotary wing”.
In the general category of rotorcraft various distinct types may be distinguished, however the hybrid helicopter in accordance with the invention includes, in addition to at least one main rotor, at least one propulsive propeller, preferably two variable-pitch propulsive propellers, forming parts of thrust units located on the left and right sides of the hybrid helicopter.
In a conventional helicopter, at least one main rotor under appropriate drive serves to provide both lift and propulsion. A helicopter is capable of hovering and of remaining stationary at a point in three dimensions, is capable of taking off or landing vertically, and can move in any direction (forwards and backwards, sideways, up and down).
Regulation of the power driving a conventional helicopter (not having any propeller thruster(s)) generally comprises a control member or module that adapts the power delivered by the power plant to the power required by the moving assemblies (rotor(s) and accessories), in such a manner as to maintain the speed of rotation of the main rotor(s) and of the power transmission system at a setpoint value.
On an aircraft propelled by one or more propellers of variable pitch, power regulation generally includes a regulator member and/or module (generally of the hydromechanical type) that adapts the pitch of the propulsive propeller so as to consume all of the available power that results from the pilot operating a throttle (or thrust) control member or lever.
Those two types of regulation cannot be juxtaposed in order to regulate the power of a rotorcraft fitted with propeller thruster(s), since the regulation techniques are in opposition. The member for adapting power during conventional regulation of a helicopter would counter transient speed variation in the power transmission system that would result from causing thrust delivered by the propeller(s) to vary.
Furthermore, for a rotorcraft fitted with propeller thruster(s), regulation of said propeller thruster(s) whereby the pilot varies propulsion propeller pitch directly might give rise to propeller damage as a result of sudden variations in the driving torque delivered thereto.
In a hybrid helicopter having variable pitch propellers, the pilot must simultaneously limit the variation of the collective pitch in an upward direction and consequently the variation in the power transmitted to the rotor(s) by the power plant via the transmission members in order to avoid exceeding the mechanical or thermal limits of said elements, and also, and for the same reasons, limit the thrust control applied to the propellers, i.e. the propeller pitch control, as explained above.
Furthermore, since the power plant of a hybrid helicopter is constituted by one or more turbine engines, the speeds of rotation of the outlet from the turbine engine(s), of the propeller(s), of the rotor(s), and of the mechanical system interconnecting them are mutually proportional, with the proportionality ratio being constant regardless of the flight configuration of the hybrid helicopter under normal conditions of operation of the integrated drive train.
It can thus be understood that if a hybrid helicopter is fitted with a single turbine engine, that engine rotates the rotor(s) and the propeller(s) via the mechanical interconnection system. However, if two or more turbine engines are fitted to the hybrid helicopter, then the rotor(s) and the propeller(s) are driven in rotation via the mechanical interconnection system by said turbine engines.
In other words, the drive train operates without any option for varying the ratios of the speeds of rotation between the turbine engine(s), the propeller(s), the rotor(s), and the mechanical interconnection system.
Consequently, the rotor(s) is/are advantageously always driven in rotation by the turbine engine(s) in normal flight configurations, and it/they always develop(s) lift, whatever the configuration of the aircraft.
More precisely, the rotor(s) is/are thus designed to provide the hybrid helicopter with all of its lift during stages of take-off, landing, and vertical flight, and part of its lift during cruising flight, with an auxiliary wing then contributing part of the lift for supporting said hybrid helicopter.
Thus, the rotor(s) provide(s) some of the lift of a hybrid helicopter in cruising flight together optionally with a small contribution to the propulsion or traction forces (in a helicopter), but without any contribution to drag. These operating conditions thus lead to a reduced amount of power dedicated to providing traction being delivered to the rotor(s). A small contribution to the propulsion forces is obtained by a small amount of inclination of the rotor disk(s) towards the front of the aircraft. This process degrades the lift-to-drag ratio of the rotor(s) very little and consequently it is more advantageous in terms of power balance than is a request for additional thrust to be exerted by the propeller(s).
Advantageously, the wing comprises two half-wings, each half-wing being on a respective side of the fuselage. The half-wings may together make up a high wing, in which case they preferably present a negative dihedral angle. Nevertheless, they may also constitute either a low wing, preferably with a positive dihedral angle, or indeed an intermediate wing with any dihedral angle. Depending on the variant, the plane shape of these half-wings may correspond to rectangular half-wings, tapering half-wings, forwardly or rearwardly swept wings, etc.
Below, reference is made to a movable aircraft tailplane or motor-driven tailplane, meaning that the deflection angle of said tailplane may take different values.
Furthermore, the term “aircraft” is used quite generically for whatever kind of aircraft constitutes the topic under discussion.
In an “airplane” configuration, the tailplane is the elevator control of the aircraft. It is a member for directing controlling pitch and vertical speed of the aircraft.
In a “helicopter” configuration, the tailplane of the aircraft is usually stationary. A tailplane with a variable deflection angle on a helicopter may be used to counter the “attitude hump” phenomenon or to fly with a level attitude in forward flight so as to diminish drag or so as to diminish the nose-up attitude when the center of gravity of the helicopter is offset a long way towards the rear. Nevertheless, as from a certain speed of advance, a conventional helicopter cannot retain a level attitude over its entire centering range (i.e. the range over which the center of gravity of the aircraft may vary) since the bending moment exerted on the rotor mast, commonly referred to as the “mast moment”, would become too great. The benefit of a tailplane with a variable deflection angle is consequently obtained at speeds of advance that are low or moderate. The above applies a priori to a tailplane that is stationary with or without flaps.
For example, document FR 2 916 420 discloses a rotorcraft having a motor-driven tailplane with motor-driven elevation control surfaces for conserving a zero mast moment in forward flight. This makes it possible to reduce forces on various mechanical parts.
A zero mast moment corresponds to a particular operating point of the aircraft and consequently allows for only very limited optimization of the power balance. Furthermore, in a certain number of flight configurations, it is not possible to conserve a zero mast moment.
In addition, a configuration that minimizes the mast moment in cruising flight is not necessarily a configuration that minimizes the power required.